The present invention relates to a method for repairing a composite-material panel of an aircraft, as well as a tool for implementing the method.
An aircraft includes numerous panels made of composite material. For example, the fuselage and the wing system of an aircraft are made up of juxtaposed panels of composite material forming the outer envelope of the aircraft.
During operation of an aircraft, some composite-material panels may become damaged, for example following an impact, and need to be repaired.
Document FR 2901246 describes a first operating mode for repairing a damaged fuselage zone comprising the steps of cutting out at least one section from the fuselage to form an opening encompassing the damaged zone and attaching, using attachment elements such as bolts or rivets, a portion of sheet metal compatible with the rest of the fuselage to close the opening.
This first operating mode is not fully satisfactory since the attachment elements remain visible and have a negative visual appearance.
According to a second operating mode intended to overcome the aforementioned drawbacks, a repair method comprises steps involving applying plies of fibers pre-impregnated with resin to the damaged zone, and covering the different plies using a tool that comprises different layers covered by a bladder or a vacuum bag.
The tool also includes a vacuum system designed to aspirate the gases present in the volume delimited by the panel and the vacuum bag. To ensure the polymerization or consolidation of the plies of fibers pre-impregnated with resin, and the adhesion of same to the rest of the panel, the panel and the tool are placed in an autoclave where the fiber plies are subjected to temperature and pressure cycles.
This second operating mode results in a near-invisible repair. However, it can only be used if the panel is removable and can be placed in an autoclave. Consequently, this second operating mode cannot be used to repair panels on the fuselage or wing system of an aircraft.
Document U.S. 2011/0067359 proposes a method for consolidating or polymerizing a composite-material panel without using an autoclave. According to this method, as illustrated in FIG. 1, a group of fiber plies 10 pre-impregnated with resin is positioned on a supporting part 12 and covered by a tool that includes different layers (not shown), a heating cover 14 and two vacuum bags: an inner vacuum bag 16 that covers the group of fiber plies 10 and an outer vacuum bag 18 that covers the inner vacuum bag 16. The inner and outer vacuum bags 16, 18 are joined sealingly to the supporting part 12 by sealing means 20 and 22.
This method can be used to repair a panel on a fuselage or a wing system of an aircraft on site. Consequently, during polymerization or consolidation of the fiber plies, the heating cover 14 makes it possible to generate the temperature cycle, and the pressurized gas injection between the two vacuum bags enables the inner vacuum bag 16 to exert a pressure on the pre-impregnated fiber plies if the atmospheric pressure exerts a pressure on the outer vacuum bag 18 that exceeds a given threshold.
In a variant of this method, a single vacuum bag is used, covering all of the fiber plies and a vacuum system enabling the gases present in the volume delimited by the vacuum bag to be removed. The atmospheric pressure then exerts a pressure on the vacuum bag that enables polymerization of the fibers under good conditions.
These repair methods cannot be used in certain circumstances where atmospheric pressure is not sufficient, for example when performing a repair at altitude.